Results for “attack modeling” — Page 16

NASA NTRS1983

Experiments in a three-dimensional adaptive-wall wind tunnel

Three dimensional adaptive-wall experiments were performed in the Ames Research Center (ARC) 25- by 13-cm indraft wind tunnel. A semispan wing model was mounted to one sidewall of a test section with solid sidewalls, and slotted top and bottom walls. The test section had separate top and bottom plenums which were divided into streamwise and cross-stream compartments. An iterative procedure was demonstrated for measuring wall interference and for adjusting the plenum compartment pressures to eliminate such interference. The experiments were conducted at a freestream Mach number of 0.60 and model angles of attack between 0 and 6 deg. Although in all the experiments wall interference was reduced after the plenum pressures were adjusted, interference could not be completely eliminated.

Schairer, E. T.↗

NASA NTRS1983

Thrust reversing effects on horizontal tail effectiveness of twin-engine fighter aircraft

The Langley Research Center has conducted an experimental program to determine the interference effects of thrust reversing on horizontal tail effectiveness of a twin engine, general research fighter model at approach (Mach number 0.15) and in-flight (Mach number 0.60 and 0.90) speeds. Twin vertical tails were tested at three longitudinal locations. Two nonaxisymmetric nozzle reverser concepts were studied. The effects of thrust reversing on horizontal tail effectiveness were found to be very dependent upon vertical tail locations. At approach speeds thrust reverser operation usually resulted in large variations in horizontal tail effectiveness as either nozzle pressure ratio or model angle of attack was varied. Either increases or decreases in tail effectiveness ocurred due to reverse operation depending upon tail location. At in-flight conditions there were always decreases in tail effectiveness due to reverser operation regardless of vertical tail location.

Capone, F. J.↗

NASA NTRS1985

Vapor-screen technique for flow visualization in the Langley Unitary Plan Wind Tunnel

The vapor-screen technique for flow visualization, as developed for the Langley Unitary Plan Wind Tunnel, is described with evaluations of light sources and photographic equipment. Test parameters including dew point, pressure, and temperature were varied to determine optimum conditions for obtaining high-quality vapor-screen photographs. The investigation was conducted in the supersonic speed range for Mach numbers from 1.47 to 4.63 at model angles of attack up to 35 deg. Vapor-screen photographs illustrating various flow patterns are presented for several missile and aircraft configurations. Examples of vapor-screen results that have contributed to the understanding of complex flow fields and provided a basis for the development of theoretical codes are presented with reference to other research.

Morris, O. A.↗

NASA NTRS1992

Aircraft control at high-alpha by tangential blowing

An experimental study has been undertaken to investigate the possibility of using tangential blowing to provide control forces and moments on a delta wing-forebody combination at high angles of attack. The present model can utilize the blowing separately on the wing and forebody. This research emphasizes the effect of blowing on the aerodynamics of the state of the art model by using: (1) flow visualization (smoke-flow visualization using a laser sheet), and (2) force and moment measurements. Experiments revealed that the effect of the blowing on a body-wing combination differs in several respects compared to that of wing alone or body alone. Force and moment measurements on the present model showed that the asymmetric blowing from the fuselage creates larger side force, rolling and yawing moments compared to blowing from the wing. In the presence of the forebody, the effectiveness of the rolling moment generated by blowing from the wing is reduced.

Celik, Zeki Z.↗

NASA NTRS1992

Heat flux and shock shape measurements on an Aeroassist Flight Experiment model in a high enthalpy free piston shock tunnel

This paper describes measurements of heat flux and shock shapes made on a 2.08 percent scale model of the proposed Aeroassist Flight Experiment model in a high enthalpy free piston shock tunnel T3 at the Australian National University in Canberra, Australia. The enthalpy and Reynolds number range covered were 7.5 MJ/kg to 20 MJ/kg and 150,000 to 270,000 per meter respectively. The test Mach number varied between 7.5 and 8. Two test gases, air and nitrogen, were used and the model angle of attack varied from -10 deg to +10 deg to the free stream. The results are discussed and compared to the Mach 10 cold hypersonic air data as obtained in the Langley 31 inch Mach 10 Facility as well as the perfect gas CFD calculations of NASA LaRC.

Gai, S. L.↗

NASA NTRS1992

Pitot survey of exhaust flow field of a 2-D scramjet nozzle at Mach 6 with air or freon and argon used for exhaust simulation

A pitot-rake survey of the simulated exhaust of a half-span scramjet nozzle model was conducted in the Langley 20-Inch Mach 6 Tunnel to provide an additional data set for computational fluid dynamics (CFD) code comparisons. A wind-tunnel model was tested with a 26-tube pitot rake that could be manually positioned along the mid-semispan plane of the model. The model configuration had an external expansion surface of 20 degrees and an internal cowl expansion of 12 degrees; tests were also performed with a flow fence. Tests were conducted at a free-stream Reynolds number of approximately 6.5 x 10(exp 6) per foot and a model angle of attack of -0.75 degrees. The two exhaust gas mediums that were tested were air and a Freon 12-argon mixture. Each medium was tested at two jet total pressures at approximately 28 and 14 psia. This document presents the flow-field survey results in graphical as well as tabular form, and several observations concerning the results are discussed. The surveys reveal the major expected flow-field characteristics for each test configuration. For a 50-percent freon 12 and 50-percent argon mixture by volume (Fr-Ar), the exhaust jet pressures were slightly higher than those for air. The addition of a flow fence slightly raised the pitot pressure for the Fr-Ar mixture, but it produced little change for air. For the Fr-Ar exhaust, the plume was larger and the region between the shock wave and plume was smaller.

Monta, William J.↗

NASA NTRS1958

Exploratory Wind-Tunnel Investigation to Determine the Lift Effects of Blowing over Flaps from Nacelles Mounted Above the Wing

An exploratory wind-tunnel investigation has been made to determine the lift effects of blowing from nacelles over the upper surface of flaps on a model having a delta wing of aspect ratio 3. Several flap conditions were examined. High-pressure air was blown from an external-pipe arrangement supported above the wing to simulate jet-engine exhaust. The jet momentum- coefficient range was from 0 to 3.0 and the model angle of attack was 0 deg. The results of this limited investigation show that values of jet circulation lift coefficient larger than the Jet reaction were produced with blowing over flaps from nacelles mounted above the wing. 'I!heuse of double slotted flaps with the gap unsealed between the flaps and wing had a large detrimental effect on the lift capabilities. With these gaps sealed, larger lift coefficients were obtained when fantails were added to the nacelles. The longitudinal trim problems created by large diving moments were similar to those encountered with other jet-augmented-flap systems

Riebe, John M↗

NASA NTRS1992

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot transonic wind tunnel, volume 2

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Marroquin, J.↗

NASA NTRS1992

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot Transonic wind tunnel (IA613A), volume 1

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Marroquin, J.↗

NASA NTRS1993

Transitional flight characteristics of a geometrically simplified STOVL model

The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration were measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment was designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important powered lift flow parameters. The model consists of a 60 deg cropped delta wing planform; a blended fuselage; and tandem, circular, high pressure air jets that exit perpendicular to the flat lower surface. Freestream Mach number is limited to a maximum 0.2. Model angle of attack ranges from -10 deg to +20 deg. The nozzle pressure ratios of both jets are varied between 1 and 3, and the jet exit temperatures are maintained at near ambient conditions. Detailed surface pressure measurements show that suction pressure peaks located on the upper surface of the wing during conventional wingborne flight for angles of attack greater than 5 deg move to the wing lower surface at angles of attack less than 0 deg. A reduction in these suction pressure peaks is observed when the lift jets are operating. With sonic jet exit conditions, a freestream Mach number of 0.14, and 0 deg angle of attack, the jet-induced suckdown is equivalent to a 3.7 deg reduction in angle of attack. Schlieren, laser light sheet flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.

Roth, Karlin R.↗

NASA NTRS1994

Experimental studies of transonic flow field near a longitudinally slotted wind tunnel wall

The results of detailed parametric experiments are presented for the near-wall flow field of a longitudinally slotted transonic wind tunnel. Existing data are reevaluated and new data obtained in the Langley 6- by 19-inch Transonic Wind Tunnel are presented and analyzed. In the experiments, researchers systematically investigate many pertinent wall-geometry variables such as the wall openness and the number of slots along with the free stream Mach number and model angle of attack. Flow field surveys on the plane passing through the centerline of the slot were conducted and are presented. The effects of viscosity on the slot flow are considered in the analysis. The present experiments, combined with those of previous investigations, give a more complete physical characterization of the flow near and through the slotted wall of a transonic wind tunnel.

Everhart, Joel L.↗

NASA NTRS1955

Investigation in the Langley Free-Flight Tunnel of the Low-Speed Stability and Control Characteristics of a 1/10-Scale Model Simulating the Convair F-102A Airplane

An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model simulating the Convair F-102A airplane has been made in the Langley free-flight tunnel. The model in its basic configuration and with two modifications involving leading- edge slats and an increase in vertical-tail size was flown through a lift-coefficient range from 0.7 to the stall. Only relatively low-altitude conditions were simulated. The longitudinal stability characteristics of the model were considered satisfactory for all conditions investigated. The lateral stability characteristics were considered satisfactory for the basic configuration over the lift-coefficient range investigated, except near the stall, where large values of static directional instability caused the model to be directionally divergent. An 80-percent increase in vertical-tail area increased the angle of attack at which the model became directionally divergent. The longitudinal and lateral control characteristics were generally satisfactory. Although the adverse sideslip characteristics for the model were considered acceptable over the angle-of-attack range, analysis indicates that the adverse sideslip characteristics of the airplane may be objectionable at high angles of attack.

Boisseau, Peter C.↗

NASA NTRS2006

Planar laser-induced fluorescence (PLIF) investigation of hypersonic flowfields in a Mach 10 wind tunnel

Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize four different hypersonic flowfields in the NASA Langley Research Center 31-Inch Mach 10 Air wind tunnel. The four configurations were: (1) the wake flowfield of a fuselage-only X-33 lifting body, (2) flow over a flat plate containing a rectangular cavity, (3) flow over a 70deg blunted cone with a cylindrical afterbody, formerly studied by an AGARD working group, and (4) an Apollo-geometry entry capsule - relevant to the Crew Exploration Vehicle currently being developed by NASA. In all cases, NO was seeded into the flowfield through tubes inside or attached to the model sting and strut. PLIF was used to visualize the NO in the flowfield. In some cases pure NO was seeded into the flow while in other cases a 5% NO, 95% N2 mix was injected. Several parameters were varied including seeding method and location, seeding mass flow rate, model angle of attack and tunnel stagnation pressure, which varies the unit Reynolds number. The location of the laser sheet was as also varied to provide three dimensional flow information. Virtual Diagnostics Interface (ViDI) technology developed at NASA Langley was used to visualize the data sets in post processing. The measurements demonstrate some of the capabilities of the PLIF method for studying hypersonic flows.

Danehy, Paul M.↗

NASA NTRS2020

FLEET Velocimetry in the Common Research Model’s Wing Wake

Femtosecond laser electronic excitation tagging was used to make velocity measurements in the wake of the wing of the Common Research Model (CRM). Experiments were performed in the NASA Langley Research Center’s National Transonic Facility over a range of tunnel operating conditions. Pressures ranged from 205 to 411 kPa, temperatures from 278 to 323 K, and Mach from 0.1 to 0.9. Velocity was also determined over a range of model angles of attack. Time-averaged velocity results were obtained in both air and nitrogen, while single-shot velocity was obtained under certain tunnel operating conditions. Spatially-resolved, two dimensional, single component velocity measurements were achieved using a newly-developed laser scanning technique, which proved sufficiently sensitive to measure an approximately 5% velocity deficit in the wing wake region of the CRM.

Daniel T Reese↗

NASA NTRSDate not supplied

Implementation of an Unsteady PSP System in the NASA Transonic Dynamics Tunnel

An unsteady pressure-sensitive paint (uPSP) system has been developed to provide time-resolved pressure measurements in the NASA Langley Transonic Dynamics Tunnel (TDT). Obtaining these measurements necessitated the development of environmental enclosures to protect the high-speed camera and ultraviolet lights required for uPSP from the harsh environment present during tunnel operation. Since the facility main drive was non-functioning during the testing window, performance of the uPSP system was demonstrated using an impinging jet with a passive oscillator attachment to provide unsteady flow with a known frequency independent of amplitude. Measurements were obtained for tunnel pressures ranging from 565 to 2116 psf, and model angles of attack between -4 degrees and 4 degrees. Results indicate that the system is capable of measuring surface pressure differentials on the order of 0.01 psi at full scale with a camera frame rate of at least 10 kHz. Spectral analysis shows that the fundamental frequency of the oscillating jet is captured by the uPSP system, as are the second and third harmonics. Dynamic mode decomposition highlights the dominant coherent spatial structures of the surface pressure, along with the associated frequency and growth rate of each mode, allowing for a de-noised reconstruction of the uPSP measurements. The experimental campaign outlined within this report also confirmed compatibility of the uPSP system with the TDT facility data acquisition system, and verified the successful integration with existing processing capabilities within the NASA advanced supercomputing environment.

unsteady pressure-sensitive paint↗

NASA NTRSDate not supplied

Implementation of an Unsteady PSP System in the NASA Transonic Dynamics Tunnel

An unsteady pressure-sensitive paint (uPSP) system has been developed to provide time-resolved pressure measurements in the NASA Langley Transonic Dynamics Tunnel (TDT). Obtaining these measurements necessitated the development of environmental enclosures to protect the high-speed camera and ultraviolet lights required for uPSP from the harsh environment present during tunnel operation. Since the facility main drive was non-functioning during the testing window, performance of the uPSP system was demonstrated using an impinging jet with a passive oscillator attachment to provide unsteady flow with a known frequency independent of amplitude. Measurements were obtained for tunnel pressures ranging from 565 to 2116 psf, and model angles of attack between -4 degrees and 4 degrees. Results indicate that the system is capable of measuring surface pressure differentials on the order of 0.01 psi at full scale with a camera frame rate of at least 10 kHz. Spectral analysis shows that the fundamental frequency of the oscillating jet is captured by the uPSP system, as are the second and third harmonics. Dynamic mode decomposition highlights the dominant coherent spatial structures of the surface pressure, along with the associated frequency and growth rate of each mode, allowing for a de-noised reconstruction of the uPSP measurements. The experimental campaign outlined within this report also confirmed compatibility of the uPSP system with the TDT facility data acquisition system, and verified the successful integration with existing processing capabilities within the NASA advanced supercomputing environment.

unsteady pressure-sensitive paint↗

NASA NTRSDate not supplied

Implementation of an Unsteady PSP System in the NASA TDT

An unsteady pressure-sensitive paint (uPSP) system has been developed to provide time-resolved pressure measurements in the NASA Langley Transonic Dynamics Tunnel (TDT). Obtaining these measurements necessitated the development of environmental enclosures to protect the high-speed camera and ultraviolet lights required for uPSP from the harsh environment present during tunnel operation. Since the facility main drive was non-functioning during the testing window, performance of the uPSP system was demonstrated using an impinging jet with a passive oscillator attachment to provide unsteady flow with a known frequency independent of amplitude. Measurements were obtained for tunnel pressures ranging from 565 to 2116 psf, and model angles of attack between -4 degrees and 4 degrees. Results indicate that the system is capable of measuring surface pressure differentials on the order of 0.01 psi at full scale with a camera frame rate of at least 10 kHz. Spectral analysis shows that the fundamental frequency of the oscillating jet is captured by the uPSP system, as are the second and third harmonics. Dynamic mode decomposition highlights the dominant coherent spatial structures of the surface pressure, along with the associated frequency and growth rate of each mode, allowing for a de-noised reconstruction of the uPSP measurements. The experimental campaign outlined within this report also confirmed compatibility of the uPSP system with the TDT facility data acquisition system, and verified the successful integration with existing processing capabilities within the NASA advanced supercomputing environment. Note: this presentation is an MP4 video with sound, color with a run time of 10 minutes 37 seconds.

unsteady pressure-sensitive paint↗

DOE OSTI2024

Hiding-in-Plain-Sight (HiPS) Attack on CLIP for Targetted Object Removal from Images

Machine learning models are known to be vulnerable to adversarial attacks, but prior works have mostly focused on single-modalities. With the rise of large multi-modal models (LMMs) like CLIP, which combine vision and language capabilities, new vulnerabilities have emerged. However, these multimodal targeted attacks aim to completely change the model's output to what the adversary wants. In many realistic scenarios, an adversary might seek to make only subtle modifications to the output, so that the changes go unnoticed by downstream models or even by humans. We introduce Hiding-in-Plain-Sight (HiPS) attacks, a novel class of adversarial attacks that subtly modifies model predictions by selectively concealing target object(s), as if the target object was absent from the scene. We propose two HiPS attack variants, HiPS-cls and HiPS-cap, and demonstrate their effectiveness in transferring to downstream image captioning models, such as CLIP-Cap, for targeted object removal from image captions.

Daw, Arka [ORNL] (ORCID:0009000633191271)↗